The present invention relates generally to gas turbine engines, and, more specifically, to turbines therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in turbine stages which power the compressor and a shaft that typically drives a fan in an aircraft turbofan engine application.
A high pressure turbine (HPT) directly follows the combustor and receives the hottest gases therefrom from which energy is initially extracted. A low pressure turbine (LPT) follows the HPT and extracts additional energy from the gases.
As energy is extracted from the gases in the various turbine stages, the velocity and pressure distributions correspondingly vary, which in turn requires correspondingly different aerodynamic profiles of the turbine stator vanes and rotor blades. The size of the vanes and blades typically increases in the downstream direction for providing more surface area to extract energy from the combustion gases as the pressure thereof decreases.
The velocity of the gases also decreases as energy is extracted and the flowpath area increases, which in turn leads to changes in the span and thickness aspect ratios of the vanes and blades and corresponding camber thereof.
Fundamental to turbine efficiency is the aerodynamic performance of the individual turbine airfoils as the combustion gases are split along the leading edges thereof for corresponding flow along the generally concave pressure side of the airfoil and the generally convex suction side thereof. Differential pressure is effected between the opposite airfoil sides, and aerodynamic contour or camber of the airfoil is optimized for maximizing differential pressure without undesirable flow separation of the gases over the suction side.
The turbine flowpath is defined circumferentially between adjacent airfoils as well as radially between inner and outer flowpath surfaces. For the turbine nozzle, inner and outer bands integral with the vanes bound the flow. And for the turbine blades, radially inner platforms and radially outer tip shrouds bound the combustion gases.
A particular problem affecting turbine efficiency is the generation of undesirable vortices as the combustion gases are split along the airfoil leading edges near a flow boundary, such as the radially inner blade platforms. Two horseshoe vortices flow downstream on opposite sides of each airfoil and create undesirable turbulence in the flow. This turbulence can increase platform heating. And, migration of the vortices radially outwardly can decrease turbine efficiency.
The outer and inner flowpath boundaries in the typical gas turbine engine are axisymmetrical with constant diameter or radius from the axial centerline axis of the engine. The blade platforms, for example, are therefore axisymmetric with uniform circumferential curvature from their upstream forward ends to their downstream aft ends notwithstanding any axial inclination or slope thereof.
In previous turbine developments, it is known to selectively contour the flowpath boundaries to minimize the adverse affects of the horseshoe vortices. However, due to the complex three dimensional (3D) configuration of the turbine stages and the correspondingly complex 3D distributions of the velocity, pressure, and temperature of the combustion gases contouring of the flowpath boundaries is equally complex and is directly affected by the specific design of the specific turbine stage.
Accordingly, known flowpath contouring is highly specific to specific turbine stages and is not readily transferable to different stages whose efficiency and performance could instead be degraded.
Adding to the complexity of turbine blade design is the need to assemble individual blades into a supporting rotor disk. Each blade typically includes an axial entry dovetail integrally joined to the platform in a unitary assembly with the airfoil. The dovetail is axially straight and is inserted axially through a corresponding axial dovetail slot in the rotor disk.
The individual platforms have axially straight circumferential edges which adjoin each other in the full row of blades. Spline seals are mounted between the platform edges to improve turbine efficiency.
However, due to manufacturing tolerances of the outer surfaces, adjacent platforms may not be fully flush after assembly. One platform may be radially higher or radially lower than the adjacent platform causing a corresponding down step or up step.
The up step can cause a substantial reduction in aerodynamic performance as the combustion gas flow is locally blocked and diverted over the step onto the next adjacent platform.
Accordingly, it is desired to provide a platform having an improved configuration for improving turbine performance and efficiency.